ULA Briefing to National Research Council. In-Space Propulsion Roadmap. March 22, Bernard Kutter. Manager Advanced Programs. File no.

Transcription

1 ULA Briefing to National Research Council In-Space Propulsion Roadmap March 22, 2011 Bernard Kutter Manager Advanced Programs File no. Copyright 2011 United Launch Alliance, LLC. All Rights Reserved.

2 Key Transportation Technologies 3 most critical transportation technologies enabling beyond LEO Exploration: Integrated Cryogenic Propulsion Stage (CPS) design Efficient cryogenic storage Cryogenic fluid transfer Lack of these 3 technologies negatively impacted Constellation The time between Ares V and I launches was reduced from 3 months (ESAS) to 1 day This short interval was the single largest risk to mission success Use of a half full EDS for the Earth departure burn increased the initial mass in LEO (IMLEO) by >30% compared to a full EDS On-orbit fueling and efficient CPS IMLEO could be reduced by >35%. Orion's propulsion module was switched from LO2/LH2 to LO2/LCH and finally storable propulsion The resulting increased lift requirement rippled through the entire launch system forcing significant redesign File no. 1

3 CPS In-Space Applications Satellite Launch Lunar Landers Orion Service Module Propellant Depot Propellant Tanker Multi-Year CPS MMSEV CPS Courtesy NASA CPS Spans Entire Space Transportation Architecture File no. 2

4 8 7 Integrated CPS Design is Critical Centaur Derived (MF=0.; ISP=60) Full EDS (MF=0.85; ISP=8) Example 100 mt Propellant Load Possible (MF=0.5; ISP=68) Propellant Load: 100 mt Efficient in-space stage design can nearly triple delivered payload High ISP LO2/LH2 propulsion Fuel stage on orbit Efficient cryo storage Integrated stage design LEO to L2 to LS V (km/sec) 6 5 LEO to LS LEO to Phobos (propulsive) LEO to TMI LEO to LLO LEO to L2 Mars Example Lunar Example 3 LEO to TLI Ares V EDS Derived (MF=0.7; ISP=8) Payload Mass (mt) CPS Design Drives Launch Requirements File no. 3

5 HEFT Assumed CPS Inadequate CFM Technology Adversely Drives CPS Design File no.

6 CPS Design Differences CPS design has huge impact on performance Both stage mass and cryogenic storage Centaur Delta 6 klb 6 klb LO2 & LH2 5 klb 7 klb Stage Dry Mass 2 klb klb Structure 0% 87% Mass Fraction Centaur m Delta IV Integrated CPS Design Improves System Capability File no. 5

7 Mass Fraction > 0.0 Long duration Mission Flexibility Vehicle Equipment on Thermally Isolated Shelf 1 to RL-10 or NGE Class Engines File no. 6 Integrated CPS Design Advanced Common Evolved Stage (ACES) Monocoque CRES Tank with advanced common bulkhead 5 m diameter MLI enshrouded LO2 Tank Integrated Vehicle Fluids LH2 Tank Vapor Cooled Structures Payload Interface Cryo-Upper Stage Operating Time (ksec) ULA Cryo Experience Centaur Delta Integration of Individual Technologies into Effective System is Technical Challenge ULA Ariane

8 Integrated Vehicle Fluids (IVF) Utilize Hydrogen and Oxygen to replace: Hydrazine for attitude control Helium for pressurization Large Vehicle Batteries Power Provides mission flexibility Unlimited Tank Pressurization Cycles Numerous Main Engine Burns Reaction control for attitude and translation Long mission durations Eases stage refueling and reuse Tank Pressurization Controls Attitude Thrusters Battery Electrical Generator LH2 Tank HP GO2 Internal Combustion Engine HP GH2 H2 Engine Bleed LO2 Tank H2/O2 Thruster Drive Motor O2 Pump Settling Thruster O2 & H2 Vaporizer/Heater Drive Motor H2 Pump File no. 7 United Launch Alliance (ULA) Proprietary Information

9 Mission Architectures All exploration missions require multi-launch aggregation 2 or more launches Transfer/Assembly of: Payloads, CPS, and/or propellant Multi months loiter Dual Launch L1 Gateway Mission File no. 8 HEFT NEO Mission Courtesy NASA Multi-Launch Requires Long Duration Cryo Storage and Significantly Benefits from Cryogenic Propellant Transfer

10 Long Duration CPS Earth Departure Stage File no. Mars or NEO return stage Lunar lander Propellant depot Return stage mission duration Multi year mission with very low boil-off 1 year: 0.027%/day 2 year: 0.01%/day 3 year: 0.00%/day year: 0.007%/day On orbit fueling allows: Structure/insulation to not be driven by launch environment Reduced structural heat leak paths Very high mass fraction (>0.0) LH2 Module Mission Module Centaur LO2 Module 70 mt Propellant 100 mt Pld Mars to TEI

11 Cryo Storage Experience Cryo Dewar Experience: small scale, heavy, very efficient Centaur Experience: large scale, light weight, modest efficiency TC-15 (LO2) (LH2) TC-11 (LO2) (LH2) Tank Heating (Btu/hr) Boil-Off (%/day) System B-O (%/day) Spitzer 0.05%/day boil-off Solid He Courtesy NASA/JPL-Caltech AV-005 Debris Shield AV-003 White Decal TC MLI AV-007 bare fixed foam Hydrogen Thermal Test Article 0.022%/day boil-off LH2 Courtesy NASA File no. 10 COBE 0.07%/day boil-off SfHe Courtesy NASA GSFC Need to Combine Dewar and CPS Technology to Enable Efficient, Light Weight Cryo Storage

12 Integrated Cryo Test Integrated ground cryo test Demonstrate large scale, flight like systems Use actual Centaur flight tank Demonstrate low boil-off storage ~2%/day current flight demonstrated ~0.25%/day with existing Centaur Guide future vehicle design to support <0.1%/day boil-off Enhanced Thermal Protection Vapor cool Thick MLI Courtesy NASA Vapor cool Broad area cool File no. 11 Vapor cool engine mount Technology Advancement Through Large Scale Ground Demonstration Clean bulkhead - Min penetrations

13 CRYogenic Orbital TEstbed In-space laboratory for cryo fluid management (CFM) technologies Uses residual Centaur LH2 after primary payload separation 2010 Ground Test Flight Article Design Courtesy Ball Small Scale Demonstrations Leading to Large Scale Cryo-Sat Flagship Technology Demonstrations 2015 Courtesy NASA Repeated Flight Opportunities Enabling Technology Advancement File no. 12

14 CFM Technologies TRL Cryo Fluid Management Technology Current TRL TRL Post-CRYOTE Lite TRL Post-CRYOTE Pup, Free Flier 0-g Stld 0-g Stld 0-g 10 - g Transfer System Operation 5 Pressure Control 6 Low Acceleration Settling N/A N/A N/A Tank fill operation 5 Thermodynamic Vent System Multi-layer insulation (MLI) Integrated MLI (MMOD) 6(2) 6(2) (7) (7) Vapor Cooling (H 2 para-ortho) () () Passive Broad Area Cooling (active) () () () () Active cooling (20k) Ullage and Liquid Stratification 3 Propellant acquisition 2 Mass Gauging 3 Propellant Expulsion Efficiency 3 System Chilldown 5 Subcooling P>1atm (P<1atm) (5) (5) (5) (5) (5) (5) Fluid Coupling File no. 13

15 In-Space Engine Development In-Space propulsion requirements Reliable Producible Affordable High ISP (>60 sec) Light weight (~500 lb) ~25 klb thrust Low net positive suction pressure Engine out Courtesy PWR Courtesy Xcor Continuous US Propulsion Investment File no. 1

16 Solar Electric Propulsion Solar electric propulsion has potential to significantly reduce required launch mass Typically large exploration class missions assume high power SEP 50kW class vehicles such as FTD1 Ultimately 200 kw to multi MW class At low mission tempo SEP cost may not be worth reduced launch mass Smaller SEP systems have broad application xclass robotic exploration Rideshare orbit delivery Propellant scavenging and delivery to HEO 5kW class vehicles such as ESPA OMS Courtesy NASA Courtesy Busek File no. 15 Small SEP provides valuable experience

17 Summary Enhanced technologies supporting CPS design critical for Exploration Integrated CPS design Efficient cryogenic storage Cryogenic fluid transfer Integrated Vehicle Fluids Mission capability, reliability Integrated testing Ground testing Affordable in-space testing (CRYOTE) Continuous engine investment Affordable solar electric propulsion File no. 16